CN215949589U

CN215949589U - Aeroengine blade and aeroengine comprising same

Info

Publication number: CN215949589U
Application number: CN202121362946.4U
Authority: CN
Inventors: 黄正斌; 张宝岭; 吴明峰; 邱冬梅
Original assignee: AECC Commercial Aircraft Engine Co Ltd
Current assignee: AECC Commercial Aircraft Engine Co Ltd
Priority date: 2021-06-18
Filing date: 2021-06-18
Publication date: 2022-03-04
Anticipated expiration: 2031-06-18

Abstract

The utility model provides an aircraft engine blade and an aircraft engine comprising the same, wherein the aircraft engine blade comprises a blade back and a trailing edge, a plurality of grooves are arranged at the corner where the trailing edge is located or on the blade back, the grooves are arranged along the length direction of the trailing edge, and at least one part of the grooves leaks out of the blade back. The blade back and the corner of the tail edge are provided with the grooves, the grooves are arranged on the corner of the tail edge, at least one part of the grooves is leaked out of the blade back, the grooves do not need to process the tail edge of the blade, the problem that the grooves are inconvenient to process the tail edge of the thin blade can be avoided, stress concentration of the blade body can be effectively avoided, and influence on working performance of the blade can be avoided. The grooves are arranged along the length direction of the tail edge, so that the abrasion condition of the blade is marked by utilizing the abrasion of the grooves in the same direction. The groove can visually judge the abrasion degree of each blade.

Description

Aeroengine blade and aeroengine comprising same

Technical Field

The utility model relates to an aircraft engine blade and an aircraft engine comprising the same.

Background

In rotating machinery such as an aircraft engine, a gas turbine, a steam turbine and the like, a small radial distance between the top end of a rotor blade and a stator casing is called as a blade tip clearance, and is a basic parameter in the design process of the aircraft engine, and the size of the blade tip clearance has a large influence on the performance and the structural safety of the aircraft engine. The smaller the blade tip clearance is, the lower the fuel consumption rate of the aero-engine is, the longer the service life is, and meanwhile, the performance of the aero-engine can be effectively improved. The larger the blade tip clearance is, the lower the propelling efficiency of the aeroengine is, so that the fuel consumption rate is increased due to the need of supplementing fuel, the EGT temperature is increased, and the service life of the engine is further shortened.

In the operation process of an aircraft engine, due to the problems of thermal deformation incompatibility between a rotor blade and a stator casing and the like, the honeycomb or the coating at the position corresponding to the rotor blade tip and the casing is frequently abraded, and further the blade tip is easily slightly abraded. Along with the increase of the number of the operating cycles of the aero-engine, the abrasion of the blades caused by collision and abrasion is accumulated continuously, so that the blade tip clearance of the rotor blade is enlarged, the operating efficiency of the aero-engine is reduced, the EGT temperature is increased, the service life of high-temperature parts of the aero-engine is reduced, the oil consumption rate of the aero-engine is increased, and the cost increase of the aero-engine in the operating process is greatly influenced. Because the abrasion of the rotor blade or the increase of the blade tip clearance of the engine and the performance degradation of the engine are in positive correlation, the aircraft engine generally evaluates the performance degradation condition and the health condition of the in-service engine by checking the abrasion condition or the blade tip clearance of the rotor blade in the service operation process, so as to set the maintenance level of the engine, and ensure that the operation cost and the safety of the engine are maintained within a reasonable range.

However, as aircraft engines are designed to be more compact and one engine has thousands of blades, it is difficult to measure and analyze the wear amount of the rotor blade or the blade tip clearance of the engine in service during service maintenance, and the measurement must be performed by means of complex and expensive tooling equipment, so that the operating environment is limited, and the time, labor and cost are high. Meanwhile, the establishment of maintenance schemes for complete engines or parts of in-service engines is usually determined based on the performance degradation state of the engine, the structural damage condition of the engine and the like, and the rationality and effectiveness of the establishment of the maintenance schemes are also limited by the measurement of blade tip clearances.

SUMMERY OF THE UTILITY MODEL

The utility model aims to solve the technical problem that the blade tip clearance is difficult to detect and evaluate in the prior art, and provides an aircraft engine blade and an aircraft engine comprising the same.

The utility model solves the technical problems through the following technical scheme:

the utility model provides an aeroengine blade and contain its aeroengine, aeroengine blade includes back of the leaf and trailing edge arris, and a plurality of recesses are established the corner at trailing edge arris place perhaps on the back of the leaf, it is a plurality of the recess is followed the length direction of trailing edge arris sets up, at least partly of recess leak in the back of the leaf.

In the technical scheme, the grooves are formed in the corners of the blade back and the tail edge, the grooves are formed in the corners of the tail edge, at least one part of each groove is exposed out of the blade back, and when hole detection is carried out in a narrow engine, the hole probe has good accessibility relative to the tail edge, so that the grooves can be observed in the front side conveniently, and the real accuracy of evaluation data is improved. The groove is formed in the blade back, the blade body at the tail edge of the blade is provided with the groove, the groove does not need to be machined on the tail edge of the blade, the problem that the groove is inconvenient to machine on the thin tail edge of the blade can be avoided, stress concentration of the blade body and influence on the profile of the blade body can be effectively avoided, and influence on working performance of the blade is avoided. The grooves are arranged along the length direction of the tail edge, so that the abrasion condition of the blade is marked by utilizing the abrasion of the grooves in the same direction. The groove can visually judge the wear degree of each blade without additional tooling equipment, and meanwhile, the relationship between the health states of the engine is reflected on the basis of the wear depth of the groove on the blade back of the engine, the performance decline degree of the engine, the operation cost of the engine and the like, so that a comprehensive influence factor of the wear degree of the groove on the health state of the engine is formed. The size of the blade tip gap is measured by effectively utilizing the groove, so that the health state of the engine is conveniently and visually evaluated, and the influence of a newly added structure on the aerodynamic performance of the blade body is also avoided.

Preferably, the groove is arranged on the blade back, the groove is formed on the surface of the blade back, and the groove is arranged in a non-penetrating mode.

In this technical scheme, the recess is not the penetrability structure, can not lead to blade gas leakage, can not exert an influence to blade structure and engine fuselage dynamic behavior, and structural processing also does not have the influence to its blade performance.

Preferably, the groove is disposed on the corner, and the groove is disposed on the corner in a penetrating or non-penetrating manner.

In the technical scheme, the grooves penetrate through the corners of the tail edge edges, so that the whole grooves can be detected, and the detection accuracy is improved. The grooves are arranged on the tail edge in a non-penetrating mode, so that the machining is convenient, and interference to other parts is not easy to cause during machining.

Preferably, the corner includes a trailing edge, a trailing edge face, and an abutment of the blade back and the trailing edge, and at least a portion of the groove is located in one or more of the trailing edge, the trailing edge face, and the abutment of the blade back and the trailing edge.

In the technical scheme, the grooves are formed in the tail edge, and when hole detection is carried out in a narrow engine, the hole probe has good accessibility relative to the tail edge, so that the grooves can be observed in the front side conveniently, and the real accuracy of evaluation data is improved. Set up the recess on the blade body at blade trailing edge arris department, this kind of recess need not be to blade trailing edge arris processing recess, can avoid the inconvenient problem of thinner blade trailing edge processing recess, also can effectively avoid the stress concentration of blade body and to the influence of blade body profile moreover, more can not exert an influence to blade working property. And a groove is processed on the tail edge surface, so that the cavity where the blade back is located is prevented from being damaged, and the influence on the blade of the aircraft engine is reduced.

Preferably, the groove extends through the trailing edge face.

Among this technical scheme, aeroengine blade's the wearing and tearing condition is different, probably has different wearing and tearing conditions, and the recess runs through the trailing edge face for the condition of the wearing and tearing of recess all can be detected to a plurality of angles, has increased the detection effect to the recess.

Preferably, the surface shape of the groove is an arc surface.

In the technical scheme, the cambered surface groove can prevent the tail edge from generating stress concentration and other defects, and the influence on the airflow surface of the blade body after the blade body is processed into the blade wear amount marking structure is prevented.

Preferably, the surface of the groove is a mirror surface.

According to the technical scheme, the groove bottom of the groove is machined to be high in finish degree, the mirror surface effect of the groove bottom is improved, hole detection is carried out in a dark engine flow channel, a hole detector can be used for finding and accurately evaluating the blade abrasion loss more efficiently, the vertical horse and other blade body base parts are distinguished, the work load is reduced, and the accuracy of blade body abrasion loss evaluation is improved.

Preferably, the number of the grooves is three.

In the technical scheme, the number of the grooves is three, and the clearance between the blade tip and the casing of the blade can be measured by the three grooves. The interlobe clearance value is optimal when the groove closest to the casing is visible, and engine performance and health are best. When only two grooves can be seen, the blade clearance is in a reasonable range, and the influence on the performance decline and the health state of the engine is small. When the middle groove is not seen, the performance of the engine is degraded and the health state of the engine is reduced to the warning state, the attention range between the first groove and the second groove can be defined, when the blade tip clearance is within the attention range, the normal operation of the engine can cause the oil consumption of the engine to be increased to a certain extent, but the operation of the engine is still in a safe and reliable state, and only daily maintenance and inspection are needed regularly. When the last groove is not seen, the normal working state of the engine is influenced to a certain extent by the increase of the blade tip clearance, certain risk exists when the engine runs in a flight profile, and major attention needs to be paid and overhaul inspection needs to be carried out within a specified number of cycles. When the three grooves cannot be seen completely, the three grooves belong to a dangerous range, the engine cannot normally meet the requirement that the engine completes a flight task due to overlarge blade tip clearance, the blades need to be scrapped and replaced immediately, and otherwise, a large danger is possibly generated.

An aircraft engine comprising the aircraft engine blade.

Among this technical scheme, aeroengine includes the aeroengine blade, can utilize the recess on the aeroengine blade to mark the health degree of engine blade to make aeroengine operation more stable.

Preferably, the aircraft engine comprises a casing, and the minimum clearance between the casing and the aircraft engine blade is a blade tip clearance.

In the technical scheme, the blade tip clearance is a basic constant in the design process of the aero-engine, the blade tip clearance has a large influence on the performance and the structural safety of the aero-engine, but the blade tip clearance is limited to a narrow space, and the measurement of the blade tip clearance is extremely difficult.

The positive progress effects of the utility model are as follows:

Drawings

FIG. 1 is a schematic view of an aircraft engine blade according to the present invention.

FIG. 2 is a schematic tip clearance view of the aircraft engine blade shown in FIG. 1 without tip wear.

FIG. 3 is a green line value tip clearance schematic of the aircraft engine blade shown in FIG. 1.

FIG. 4 is a schematic yellow-line tip clearance of the aircraft engine blade shown in FIG. 1.

FIG. 5 is a red line value tip clearance schematic of the aircraft engine blade shown in FIG. 1.

FIG. 6 is a schematic illustration of a groove in a trailing edge of the aircraft engine blade shown in FIG. 1.

FIG. 7 is a schematic illustration of a groove of a body of the aircraft engine blade shown in FIG. 1.

FIG. 8 is a schematic cross-sectional view of a groove of the aircraft engine blade shown in FIG. 7.

FIG. 9 is a schematic illustration of a bore inspection of the aircraft engine blade shown in FIG. 1.

FIG. 10 is a schematic diagram of a hole detection inspection of a groove of the aircraft engine blade shown in FIG. 1.

FIG. 11 is a cross-sectional view of an aircraft engine of the present invention.

Description of the reference numerals

Leaf back 1

Trailing edge 2

Groove 3, first groove 31, second groove 32, third groove 33

Trailing edge face 4

Case 5

Aircraft gas turbine engine 6

Fan section 7

Compressor section 8

Combustor section 9

Turbine section 10

Compressor rotor blade 11

Compressor stator casing 12

Turbine rotor blade 13

Leaf basin 14

Leading edge blade tip 15

The pores 16

Blade tip 17

Turbine stator casing 18

Inner wall surface 19

Arcuate bottom surface 20

Detailed Description

The present invention will be more clearly and completely described in the following description of preferred embodiments, taken in conjunction with the accompanying drawings.

As shown in figures 1 and 9, the utility model discloses an aircraft engine blade which comprises a blade back 1 and a trailing edge 2, wherein a plurality of grooves 3 are formed in a corner where the trailing edge 2 is located, the grooves 3 are formed along the length direction of the trailing edge 2, and at least one part of each groove 3 leaks out of the blade back 1. The corner comprises the abutment of the trailing edge 2, the trailing edge face 4, the blade back 1 and the trailing edge 2. At least a portion of the groove 3 is located in one or more of the three at the abutment of the trailing edge 2, trailing edge face 4, blade back 1 and trailing edge 2. When hole detection is carried out in a narrow engine, the hole probe has better accessibility relative to the tail edge 2, so that the groove 3 can be observed in the front direction conveniently, and the true accuracy of the evaluation data is improved.

Set up recess 3 on the blade body at blade trailing edge arris 2 departments, this kind of recess 3 need not be to blade trailing edge arris 2 processing recess 3, can avoid the inconvenient problem of thinner blade trailing edge processing recess 3, also can effectively avoid the stress of blade body to concentrate and to the influence of blade body profile moreover, more can not exert an influence to blade working property. The groove 3 is processed on the tail edge surface 4, so that damage to a cavity where the blade back 1 is located can be avoided, and the influence on the blade of the aircraft engine is reduced. The aircraft engine blade also includes a leading edge tip 15.

As shown in fig. 1 and 2, the grooves 3 are provided in the corners either penetrating or non-penetrating. The groove 3 penetrates through the corner arranged on the tail edge 2, so that the whole groove 3 can be detected, and the detection accuracy is improved. The grooves 3 are arranged on the tail edge 2 in a non-penetrating mode, so that the machining is convenient, and interference to other parts is not easy to cause during machining.

When the groove 3 is arranged at the corner in a penetrating way, the groove 3 penetrates through the tail edge surface 4. The abrasion condition of the blade of the aircraft engine is different and can have different abrasion conditions, and the groove 3 penetrates through the tail edge surface 4, so that the abrasion condition of the groove 3 can be detected at a plurality of angles, and the detection effect on the groove 3 is increased.

The aero-engine blade comprises a blade basin 14, and the groove 3 is formed in the trailing edge 2 to prevent extra increase of inflow gas from leaking from the blade basin 14 to a blade back 1 to reduce work efficiency of a compressor or a turbine. Meanwhile, the influence of a newly added structure on the aerodynamic performance of the blade body is also avoided.

In other embodiments, a plurality of grooves 3 are arranged on the blade back 1, the blade body at the tail edge of the blade is provided with the grooves 3, the grooves 3 do not need to process the tail edge edges 2 of the blade, the problem that the grooves 3 are inconvenient to process the thin tail edge of the blade can be avoided, stress concentration of the blade body and influence on the profile of the blade body can be effectively avoided, and influence on working performance of the blade is avoided.

As shown in fig. 1, the groove 3 provided on the blade back 1 is formed on the surface of the blade back 1, and the groove 3 is provided on the surface of the blade back 1 without penetrating. The groove 3 is not a penetrating structure, so that the gas leakage of the blade can not be caused, the power performance of the blade structure and the engine body can not be influenced, and the structural processing also has no influence on the performance of the blade.

As shown in fig. 1 and 6, a plurality of grooves 3 are provided along the length direction of the trailing edge 2 so that the wear condition of the vane is marked by the wear of the plurality of grooves 3 in the same direction. The groove 3 can visually judge the wear degree of each blade without additional tooling equipment, and meanwhile, the relation between the health states of the engine is reflected on the basis of the wear depth of the groove 3 on the blade back 1 of the engine, the performance decline degree of the engine, the operation cost of the engine and the like, so that a comprehensive influence factor of the wear degree of the groove 3 on the health state of the engine is formed. The size of the blade tip clearance is measured by effectively utilizing the groove 3, so that the health state of the engine is conveniently and visually evaluated, and the influence of a newly added structure on the aerodynamic performance of the blade body is also avoided.

As shown in fig. 7 and 8, the surface shape of the groove 3 is a curved surface. Cambered surface recess 3 can prevent that trailing edge department from producing stress defect such as concentrated, prevents to produce the influence to the air current profile of blade body behind blade body processing blade wearing and tearing volume mark structure, avoids producing stress defect such as concentrated to influence blade intensity.

As shown in fig. 7 and 9, the arc-shaped bottom surface 20 of the groove 3 is a mirror surface. As shown in fig. 9 and 10, the arc-shaped bottom surface 20 of the groove 3 is processed with high finish, the mirror surface effect of the arc-shaped bottom surface 20 at the groove bottom is improved, hole detection is performed in a dark engine flow passage, a hole detector can be used for finding and accurately estimating the blade abrasion loss more efficiently, a stand horse can be distinguished from other blade body base parts, the working load is reduced, and the accuracy of estimation of the blade abrasion loss of a substitute blade body is improved.

As shown in fig. 2 to 5, the number of the grooves 3 is three. The three grooves 3 are a first groove 31, a second groove 32 and a third groove 33 from the part close to the blade tip 17 to the part far away from the blade tip. The clearance between the tip of the blade and the casing 5 can be measured by three grooves 3. The distance from the blade tip to the first groove 31 is a, the distance from the second groove 32 to the blade tip is b, the distance from the blade tip to the third groove 33 is c, and different distances from the groove 3 to the blade tip 17 can be used as references for different tip wear depths of the blade tip 17.

Since the radial heights of the first groove 31, the second groove 32 and the third groove 33 are sequentially reduced, a corresponding relationship can be established between the radial positions of the first groove and the second groove and the tip clearance of the aircraft engine blade. When the engine leaves the factory, the clearance between the blade tip and the inner wall surface 19 of the casing 5 is A, and the blade tip clearance value A is in the optimal state. While the location of the first slot 31 may represent the amount of tip wear or green line value of the tip clearance, i.e., the gradual wear of the blade tip 17 over the cumulative number of engine operating cycles.

As shown in fig. 3, when the first flute 31 is worn and just cannot be seen, the corresponding green line tip clearance value is B, which can define that the tip clearance is in a reasonable range from a to B, and the influence on the performance degradation and the health condition of the engine is small and within an acceptable range. The position of the second groove 32 may represent a blade tip wear amount or a blade tip clearance, as shown in fig. 4, that is, when the blade tip 17 is worn until the second groove 32 is just not seen, it indicates that the performance of the engine is degraded and the engine is in a state of being in which the state of being in a state of being in which the engine is in which the state of being in which the blade tip clearance is not being in a state of being in which the state of being in a state of being in which the blade tip clearance being in a state of being in which the engine being in a state of being in which the blade tip being in a state of being in which the engine being in which the blade tip being in a state of being in which the engine being in a state of being in which the engine being in a state of being in which the engine being in which the blade being in a state of being in which the blade tip being in a state of being in which the blade being in a state of being in which the blade being in a state of being in which the blade being in a state of being in a blade being in which the blade being in a blade being in which the blade being in a blade being in which the blade being in a blade being in which the blade being in. The position of the third slot 33 may represent a red line value of the tip wear or the tip clearance, as shown in fig. 5, that is, when the blade tip 17 is worn until the third slot 33 is just not visible, the tip clearance value D is the maximum tip clearance value acceptable by the engine; the blade tip clearance value can be defined to belong to a warning range from C to D, the normal working state of the engine is influenced to a certain extent by the increase of the blade tip clearance, certain risk exists in the operation of the engine in a flight profile, and important attention needs to be paid and overhaul inspection needs to be carried out in a specified number of cycles; and defining that when the blade tip clearance value exceeds D, the blade tip clearance is too large, so that the engine cannot normally meet the requirement of completing the flight task of the engine, the blade needs to be scrapped and replaced immediately, and otherwise, a greater danger is possibly generated.

As shown in FIG. 11, the utility model also discloses an aircraft engine which comprises an aircraft engine blade. The aero-engine comprises an aero-engine blade, and the health degree of the aero-engine blade can be marked by using the groove 3 on the aero-engine blade, so that the aero-engine runs more stably.

As shown in fig. 11, the aircraft engine includes an aircraft gas turbine engine 6, the aircraft gas turbine engine 6 is mainly an axial flow type double or triple rotor engine composed of a fan section 7, a compressor section 8, a combustor section 9 and a turbine section 10, and a minute radial gap between a rotor blade 11 of a press machine and a stator casing 12 of the compressor, and a minute radial gap between a rotor blade 13 of the turbine machine and a stator casing 18 of the turbine machine are called a tip gap. The aircraft engine comprises blade tips 17, and apertures 16 are arranged between the blade tips 17.

Claims

1. The aircraft engine blade is characterized by comprising a blade back and a tail edge, wherein a plurality of grooves are formed in the corner where the tail edge is located or the blade back, the grooves are formed in the length direction of the tail edge, and at least one part of each groove is exposed out of the blade back.

2. The aircraft engine blade according to claim 1, wherein said groove is provided on said blade back, said groove is formed on a surface of said blade back, and said groove is provided non-penetratingly.

3. An aircraft engine blade according to claim 1, wherein the groove is provided in the corner and the groove is provided either through or non-through the corner.

4. The aircraft engine blade according to claim 1 wherein said corner comprises a trailing edge, a trailing edge face, and an abutment of said blade back and said trailing edge, and wherein at least a portion of said groove is located in one or more of said trailing edge, said trailing edge face, and an abutment of said blade back and said trailing edge.

5. The aircraft engine blade according to claim 4 wherein said groove extends through said trailing edge surface.

6. The aircraft engine blade according to claim 1 wherein the surface of said groove is contoured.

7. The aircraft engine blade according to claim 1 wherein the surface of said groove is a mirror surface.

8. The aircraft engine blade according to claim 1 wherein said number of grooves is three.

9. An aircraft engine, characterized in that it comprises an aircraft engine blade according to any one of claims 1 to 8.

10. The aircraft engine of claim 9, wherein said aircraft engine includes a case, and a minimum clearance between said case and said aircraft engine blade is a tip clearance.